Takeoff director systems for aircraft



July 15, 1969 G. R. SLEIGHT 3,455,150

TAKEOFF DIRECTOR SYSTEMS FOR AIRCRAFT Filed July l1, 1966 4 Sheets-Sheetl July 15, 1969 G. R. SLi-:lGHT

TAKEOFF DIRECTOR SYSTEMS FOR AIRCRAFT File-d July ll. 1966 4Sheets-Sheet il es? ns ATT PNEY`l July l5, 1969 G. R. SLi-:IGHT

TAKEOFF DIRECTOR SYSTEMS FOR AIRCRAFT Filed July l1, 1966 4 Sheets-Sheet5 July 15, 1969 G. R. sLElGHT 3,455,150

l TAKEOFF DIRECTOR SYSTEMS FOR AIRCRAFT Filed July 11. 196e 4sheets-sheet I l l 1 l I l l I l L@ I t? l l I l l l l i i I i m L 'Q KJ Si, /NVE'NTOR GEO/PGE' R. SLE/GH T E AORNEYS United States Patent O3,455,160 TAKEOFF DIRECTOR SYSTEMS FOR AIRCRAFT George R. Sleight,London, England, assignor to Elliott Brothers (London) Limited, London,England, a British company Y Filed July 11, 1966, Ser. No. 564,062Claims priority, application Great Britain, July 14, 1965,

Int. Cl. G01c 21/00 U.S. Cl. 73-178 3 Claims ABSTRACT OF THE DISCLOSUREThe pitch demand signal of a system for controlling an aircraft duringtakeoff is determined in parts by avelocity signal derived as a functionof a minimum velocity below which the aircraft is not permitted to climband the aircraft rate of climb. This velocity signal is fed through asumming device and acne-way device to `an integrator to produce anoutput signal. The output signal is inverted and fed back as an input tothe summing device so that if the velocity signal decreases, thefeedback signal becomes the greater and the output ofthe summationdevice is blocked by the one-way device. The output signal, then, mayincrease or remain steadybut may not decrease.

means to combine the measurementssignals lwith the con stant signals toprovide a composite demand signal representative of the `action requiredto perform the takeoff and initial climb cycle in accordance wtih apre-determined programme, and a directorindicatinginstrument to displaythe composite signal so that the pilot, by moving the controls to carryout the demands displayed by the system, can perform the requiredsequence of takeoff control operations in a substantially optimalmanner. The present invention is a modification of this takeoff directorsystem.

According to the invention an aircraft takeoff director system asdescribed in the above patent includes means capable of ensuring that asignal representative of demanded climb speed can be increased duringtakeoff of an aircraft but cannot be decreased during the .takeoff of.

the aircraft.

Said means may also be such that no upper limit is provided for saidsignal which is representative of demanded climb speed.

Said means may comprise a comparator to which said signal is arranged tobe fed, integrating means the input for which is derived from the outputof the cornparator via a one-way device, and the output of which isutilised as the effective demanded climb speed signal and is also fedback )as a further input to the comparator.

The feed back path from the integratingmeans to the comparator maycomprise an inverter and the comparator may comprise a summationamplifier, the one-way device being in the form of =a diode or similarunidirectional conductive device.

In the system described with reference to FIGURE-3 3,455,160 PatentedJuly 15, 1969 ICC of the above patent the signal representative ofdemanded climb speed (VTOC) is shown to be: VToc=(V2+K1(1'-K2) 0) V3where V2 is the initial climb out speed with one engine inoperative andV3 is the initial climb out speed with all engines operating, the rateof climb being denoted by h.

In the case of lan aircraft taking off, in the event of an enginefailure at or near the ground it is desirable that the aircraft climbsont at a speed V2 or slightly more if the thrust to weight ratio issufficiently high. In the all engines operating case the aircraft shouldbe climbed not revert back to the speed V2 or slightly above butcontinues to climb at the speed which had been yestablished at themoment of the engine failure.

Further the situation may arise where in a high speed jet aircraft thespeeds V2 and V3 are both well below the minimum drag speed for thataircraft and it is therefore desirable for the aircraft to continue toaccelerate on all engines. As the drag increases with increasing speedso will the rate of climb tend to increase, giving rise to an increasein demanded speed in the absence of a V3 limit. It the demanded speed isthus allowed to increase, then if the director signal is continuouslynulled it follows that the aircraft speed will require to 'be increasedand the aircraft will continue to accelerate. It then follows that insuch circumstances it may be preferable with this system to omit the V3limit.

Other objects will appear from the following description with referenceto the accompanying drawings, in which:

FIGURE l (not to scale) shows diagrammatically the takeoff and initialclimb of an aircraft, broken down into four sequential phases;

FIGURE 2 is a block schematic diagram of a simple b asic form oftake-off director system according to the invention;

FIGURE 3 is a block schematic diagram of a more complex form of take-offdirector system according to the invention;

FIGURE 4 is a view of the face of the directorin'dicating instrument;and f n FIGURE 5 is a block diagram illustrating the present invention.

Referring to the dnawings, FIGURE 1 is av diagram showing the take-offand initial climb of anv aircraft, carried out in accordance with apredetermined programme,

L broken down intofour sequential phases, as follows:

Phase A (pre-rotation) extends from the start of the take-off ground runof the aircraft '11,"up to the point" indicated by reference 12, withthe aircraft in the position 11a, at which the aircraft has attainedan'airs'peed' VR, scheduled for the beginning of rotation.

Phase B (rotation) consists of the progressive nose-up rotation of theaircraft with the main wheels still ont the runway, until lift-offoccurs when the aircraft has reached the position 13 on the runway, anditsattitude'isas dicated at 11b.

z' retracted, the acceleration being progressive reduced-as theaircraftapproaches the scheduled take-off climb speedy point 14 VTOC. At thispoint the aircraft has reached the and its attitude is as indicated by11d.

P hase D (initial climb) is the substantially constant-i speed climb up`-to a height of the order of 1500 feet, 'at point 15, the attitude ofthe aircraft being as indicated at 11e, which is the same as that shownat 11d.

In the system shown in FIGURE 2 the demands signalled to the pilotrequire the use of only two measuring instruments. These are anintegrating pitch rate gyroscope 16 having one output q and a secondintegrated output f q, and an airspeed sensor 17 having an output V.

FIGURE 2 also shows the other units included in the system. The otherunits are a time delay unit 108 which provides a delay T1, which ispreset; a limit constant unit 109 which provides a limit constant k1,which is preset; a takeoff climb-speed unit 20 which provides a signalVTOC, which is preset; a rotation speed unit 18 to provide a signal VR,which is preset; a rotation warning speed unit 1 9 to provide a signalVRW, which is preset; and a pitch-up demand unit 55 to provide a signalA0B, which is preset.

FIGURE 2 also shows a voltage source 23 feeding switches, respectively24 and 24a, which rare open until the oleo legs become fully extended asthe aircraft lifts off the ground. The switches are wired in series sothat no` signal is given until both legs are fully extended.

A phase advance unit 25 provides a time constant T2, andother unitscomprise a limiter 26, a shaping unit 27 and a Director on switch 28.

Further units contained in the system are amplifiers, respectively 34,35, 36, 37 and 38, and the system is completed by a three-way switch 40for the director indicator instrument 41.

The constants previously referred to which are yernployed in the systemaccording to FIGURE 2 are as follows.

T1, provided by the time delay unit i108, provides a time delay, whichmay be of the order of two seconds from the commencement of phase Bduring which the closure switches 24 and 24a will not initiate phase C.Phase B must therefore be at least T1 seconds in duration.

The limit constant k1 provided by the unit 109 is la limiting constantwhich is applied to the limiter 26 to prevent the signal from the phaseadvance unit 25 from exceeding a certain amplitude. The reason is thatthis signal-acts to demand a proportional nose-up pitch nate which isrequired not to exceed `a certain constant value during the early or anypart of phase C (pull-up) lest the aircraft incidence should approach toclose to the stall condition.

AThe takeoff climb-speed signal VTOC provided by the unit 20 correspondsto the correct initial climb-speed,

r The variablersignal V provided by the airspeed measuring device 17represents the instantaneous airspeed of the aircraft.

yThe constant VR provided by the unit 18 is the rotation speed which isscheduled for the particular aircraft.

The rotation warning speed signal VRW, provided by the unit 19 is toindicate that the aircraft speed is closely approaching rotation speed.y Y 1 lThe pitch-up demand signal AGR provided by the unit S ischosen tobe somewhat greater than the noseup change of attitude which will givelift-off.

The integrating pitch rate gyroscope needs no particular description,except to mention that integrating circuitry is provided sothat a ratesignal q and a second signal f q are available. ,Y

The shaping unit 27 is provided for the purpose of holding the directorindex so that it cannot move more than a prescribed amount lbelow theindicator zero pitch position,

pitches nose-up at a speed less in good time. It also provides dampingthroughout the` later stages of the takeoff.

FIGURE 4 shows the face of the director indicating instrument; Theinstrument itself may conveniently be constituted by a cathode ray tubehaving a calibrated scale mounted in front of tlie screen face. Othertypes of indicator may, of course, be used. As shown in FIGURE 4, thescale is of square shape. At the left-hand side of the instrument is ascale of pitch rate demand, the figures above the zero line indicating anose-up demand scaled in degrees per second and the figures below thezero line indicating a nose-down or speed increase demand scaled inknots. The response to a nose-down demand under particular conditions,i.e., by the pilot setting his controls appropriately to put the nose ofthe aircraft down, will lead to an increase in speed so that a nose-downdemand effectively connotes a speed increase demand. The right-handportion of the instru-ment face above the zero line is calibrated inpitch angle change demand in degrees and reads from zero to 15 The scalecontains a zero marker S6, and albove it is a Vmarker 57 which indicatesa pull-up pitch rate demand of 11/2 /sec., which is a predeterminedlimit Below the zero line 56 is a marker 458 which indicates an airspeedVRW which, in the example being described, is 10 knots less than VR. i

In operation, the cathode ray spot is made to oscillate rapidly acrossthe instrument dial to produce a horizontal line. Where the systemincludes means to indicate spiral rotation of the aircraft the line willtilt to left or right as the aircraft banks to one side or the other, asillustrated, for example, by the line 60, which represents the aircraftbanked about 12 to starboard.

The sequence and method of operation of the system are as follows:

(1) Power on.-When power is switched on, the threeway switch 40 moves toa first position in which the moving member 42 makes contact with aiirst contact stud 43. The integrating pitch rate gyroscope 16 is run upto speed. The signal q from the gyroscope 16 is applied over a line 44to the amplifier 38, whose gain is kq. The output signal from theamplifier is -kqq and passes over a line 45 to the director indicatinginstrument 41. The index of the indicator is immediately centered underthe action of the electrical spring ofthe gyroscope gimbal/ rotorsystem. When the rotor has run up to full speed the index moves aboutthe zero demand marker line with aircraft pitch rate.

(2) Dz'rector onf-Before or during taxiing to the takeoff 'startlngpoint the director system is switched on by closing switch 28. Theairspeed sensor signal V from the unit 17 and the rotation speed signalVR from the unit 18,

.i placed below centre by an amount corresponding to the' which arecombined in a differencing unit 48, produce a signal V-'VR which ispassed,together with the rotation warning speed signal VRW from the unit19, to the shaping unit 27, producing a signal representing a functionf( V-VR), which is fed to the amplifier 36. The function HIL-VR) isequal toV- VR except for V VRW when the value is limited to VRW-VR. Theamplifier 36 has a gain k., so that a signal kv(V- VR) is passed throughthe switch 28 and a line 49 to contact 43, through contact arm 42 andover a line 50 to the indicating instrument 41.

The total signal to the indicator 41 is now any Vai-afi the directorindex moves downwardly to the marker 58 which, with the scaling nsed forindex displacement per unit Valueof speed'dierenc'e- V-VR or V- VTOC, isdispreset value of VR-VRW, typically about 10 knots. The index movesabout this lower marker with aircraft pitching.

Phase A As the aircraft accelerates in its take-oi run, the directorindex still oscillates about this line due to pitching until theairspeed reaches VRW, after which the index moves upwardly towards thecentre of the indicator face at a rate proportional to the rate ofincrease of airspeed and reaches the centre point when V: VR at the endof the phase. Throughout this phase the demand is plainly nose-down,with a warning of the i-mminence of reaching VR as soon as the airspeedexceeds VRW. With VRW set at l0 knots below VR, and assuming the meanacceleration between these speeds to be three knots per second, theWarning time is just over three seconds. As in Director on the signal tothe director is the three-way switch 40 being still in position l.

Phase B.-A predetermined increase of incidence is now required and thisimplies an equal increase of pitch attitude when the aircraft is rollingalong a su-bstantially flat runway. Also, since the duration of rotationis only about 3 to 5 seconds, it is quite legitimate to ignore gyroscopedrift which, for a typical low grade instrument in this integratingmode, would be of the order of 1Arf/second (6/minute) or less.

As the airspeed reaches VR, at point 12 in FIGURE 1, a change of sign ofthe V-VR signal into the three-way switch 40 causes the switch to moveto position 2, in which the switch arm 42 engages a fixed contact 51.

A ASR signal from the unit 55 is combined with the integrated signal fqfrom the gyroscope 16 in a differencing unit 52 and fed to the amplifier37, which has a gain of k. The combined signal k(A0R-fq) is fed over aline 53 to the contact 51 and thence to the switchV arm 42 and over lineA50 to the indicator 41. This causes, in conjunction with the signal-kqq, the index to be displaced from centre by an amount proportional toku(40R-fq)kqq so that it flicks upwardly to demand nose-up rotation ofthe right amount MR at an initial rate Assuming that the pilot answersthisI demand by pulling the aircraft nose up to centre the index, thedemand shown by the index falls olf, and by the time the index is againsteady at the central position with the mean q at zero the aircraft willhave rotated stably upwards through the angle ABR.

When V= VR there is an adidtional operation, not in-v dicated in FIGURE2, whereby the integrating pitch rate gyroscope 16 is switched from therate mode (integrating capacitor charged with voltage representing AGRbut not in spring coil circuit) to the rate -i-f rate mode (capacitanceinl series 'with spring coil).

The maximum director index displacement may be made to correspond toabout oflpitch angle demand. With k=3ka this maximum displacement willalso correspond to 5/second pitch rate demand. So, Vtaking a typical AGRof 12, the index will move up to four-fifths of themaximum" at the startof phase B and this is also shown as a demand for 4/ second pitch rate.During this phase the aircraft continues to accelerate. `If AGR has beencorrectly chosen the aircraft will leave the runway de` cisively butwith very little vertical acceleration.`

Phase C.-After lift-offthe aircraft incidence is fu'rther increased toproduce'upward acceleration, in spite of the subsequent loss of lift dueto reduction of ground effect, the g being later reduced as the speedapproaches VTOC to give a smooth transitionto initial climb. The

and some of this margin must be regarded as an assurance in respect ofdrop of head component `of wind. Further-y f `complished in timegoverned by T2, probably of orderv more, this constant demandis held formost of the pullup, possibly for at least 5 seconds, so that the totalpullup time is minimized.

In the example being described the phase is initiated by full extensionof both port and starboard main oleo legs, which close microswitches 24and 24a when fully extended. These microswitches are wired in series.Any temporary simultaneous oleo extensions, due to unevenness in therunway, for example, are ignored and initiation is prohibited until atleast T1 seconds after the commencement of the phase B, the time delayT1 being provided by the signal from the unit 108. The threeway switch40 is now moved from position 2 to position 3 in which the movingcontact arm 42 engages a third fixed contact 54. Consequently, thek(A0R-f q) part of the phase B signal is replaced by a. new signal. Thisis made up of the signal VTOC, derived from the unit 20, combined withthe airspeed sensor signal V from the unit 17 in a diiierencer unit 47to provide a signal V-VTOC. This is amplified in amplifier 34, phaseadvanced in unit 2S to produce a resultant signal kv(T2V-l V- VToc) andfed to the limiter 26, which also receives the signal k1 from the limitconstant unit 109. The output signal from the limiter 26 is [kv(T2V+ V-VT0c)] k1. Thus the pull-up is demanded by displacing the director indexupwards by an amount proportional to [kv(T2Vl V- V'roc) l 1-kqfl Thiswill initially be postive if q k1/kq, and will cause the index again toflick above centre, demanding an increased pitch rate. With q indegrees/second and V, VTOC in knots, V in knots/second the pitch ratedemand should be of the order awa (wrm) te. =Va 112:6

comes olf its limit. When this happens the pilotis commanded to take oilthe pitch rate and this should be ac- During this phase underca'rriageretraction should occur.

Phase D.-When the undercarriage is fully retractedv and the air speed isWithin a preset small margin k2 (say 2` or 3 knots.) of VTOC the pitchrate signalv to the director is switched oi'because it is likely to makethe index too lively The three-way switch remains Vin position 3- butthe signal to the director is now [kvfTzV-l- V- V'roo) l k1 and nullingof this by centering the index will give stable takeoff climb atairspeed VTOC. I p

It will be appreciated that the phase-advanced speed requiring the pilotto center it. v

l This phase extends to perhaps 1,500 feet.

It will be appreciated that the phase-advanced speed error term TZV-l-V- VTOC, used in this and the previous phase may, in fact, be ofmore complex form.

Ejecr of thrust loss-If thrust loss occurs before liftoff nodiscontinuities in demand occur but the pull-up will automatically be ofshorter duration and the initial climb will be at a ilatter angle. Ifthe loss occurs very soon after lift-off there may again be no obviouschange in the time pattern of demand. Later losses will result in adownward movement of the index in sympathy with the reduction in V, andin the takeoff climb the resulting pilots action of pushing the stickforward to centre the index will check the reduction in airspeed whichmight otherwise occur. It will be appreciated that provision for varyingVTOC if a thrust loss occurs may readily be made by means of thrustsensors and logical units.

FIGURE 3 shows a more complex system which takes into account additionalcomponents of the aircrafts position and attitude and certain additionalconstants, all of which are desirable in a system intended for a largeaircraft. In addition to the pitch rate gyroscope 16 providing the q andfq signals and the airspeed sensor 17 to provide the signal V, there isa barometric height sensor 61 having an output h and an attitudereference system 62 (which also includes gyroscopes) having outputs 1;/(azimuth or heading angle), (pitch angle) and p (bank angle). Theattitude reference system is tted in the aircraft in any case, and ismade use of in the system according to the invention.

A unit 63 is provided by which the pilot can set in the actual all-upweight of the aircraft at the beginning of the trip. This provides asignal W. A further unit is a ap angle sensor unit 64 which provides asignal corresponding to the actual setting of the flaps. In the simplersystem of FIGURE 2, previously described, the ilap angle was assumed tobe constant. In the present system, however, account is taken of flapangle variation. This signal is combined with the signal representingthe all-up weight of the aircraft in a further unit 65 which mayecomprise a twoand/or three-dimensional cam arrangement or its electronicequivalent. The unit 65 provides the signal VR, which is fed to adifferencing :unit 66 where it is combined with the signal V from theairspeed sensor 17 to provide the signal V- VR which is fed to a limiter67 and will be referred to more particularly later. The unit 65 alsoprovides signals over lines 68 to the unit 63 which includes means todisplay signals corresponding to VR, V2 and V3, the latter two of whichwill be explained later.

A unit 69 contains the pilots noise-abate switch for the purpose ofreducing the engine speed to keep the aircrafts noise within acceptablelimits. When this switch is closed the unit 69 changes the values of theconstant k1 supplied by a unit 70, and a further constant signal k2supplied by a unit 71.

A unit 72 provides a constant kt which is equal to VR- VRW, the constantkr, being also applied to the limited 67.

Two phase-advance units, respectively 73 and 74, receive the Ir signalfrom the barometric height sensor and the signal V from the airspeedsensor 17 and deliver respective signals h and V. The action of the unit73 is substantially equivalent to differentiating and smoothing thesignal h which is akin to a rate signal. To the extent that h is greaterthan k2, VTOC is made greater than V2, but with an upper limit at V3.The object is to ensure that the take-oft' path never falls below thescheduled take olf net climb path.

Further adjustments are provided respectively by unitsV 75, 84 and 77 atkq, k, and kv. These are adjustments of overall gain in various parts ofthe system. The adjustment of kq sets the sensitivity of the system topitch rate changes and enables the amplitude of the pitch rate signalsto be adjusted to match the indicator scale. kv sets the sensitivity tochanges in airspeed, specifically V, while kosets the sensitivity tochanges in pitch angle.

The function unit 78 receives a bank angle signal qt from the attitudereference system 62 and also receives the signal V2 from the unit 65.The signal V2 is the take off safety speed below which the aircraft isnot permitted to climb, being variable with all-up weight etc., and isdetermined for specific conditions by the system.

The attitude reference system 62 provides the pitch attitude signal 0,the heading angle signal ,b and the bank angle signal g already referredto. The unit 79 provides a demanded heading angle signal i//D which iscombined with the heading angle signal xp in a diierencing unit 80 toprovide an output ip-tbn which is applied to a limiter 81. The limiter81 is controlled by a heading error limit signal k5 provided by aheading error limit unit x82. The purpose of limiting the signaltlf-111D is to limit the bank angle demand, however great is therequired change of heading. The maximum bank angle is laid down for allaircraft.

The constant k3, provided by a unit 76, will be referred to in detaillater.

A further unit 83 provides an adjustment which governs the lateralsensitivity of the indicator index.

As in the system of FIGURE 2, a voltage source 33 is provided, togetherwith the port and starboard oleo leg switches 24 and 24a, connected inseries and closed when the respective legs become fully extended.

The constant A0E is a signal representing an amount of nose-up rotationsomewhat greater than the minimum value which will produce lift-off. j

A time delay unit 85 provides a time delay signal T, the -purpose ofwhich is the same as the delay T1 of FIG- URE 2.

The constant k3 from the unit 76 is fed to a phase advanced air speederror computer 86, together with the signal from the function unit 78,representing g/ V sin qb, tan qb, the V2 and V3 signals from the unit`65, and the constant k1 from the unit 70. The constant k2 from the unit72 is combined with the signal h from the phase advance unit 73 in aditferencing unit 87 and the resultsignal ILL-kg is also fed to the unit86, together with the signal V' from the unit 74 and the constant kv.The output from unit 86, which is in the form together with the constantkcl from the unit 75, is fed to a pitch demand computer I88, which alsoreceives the rate or rate -l-J rate signal from the gyroscope 16 over aline 89 together with the constants kv and MR and the time delay signalT from the unit 85. Operation of the computer 88 in phase C is initiatedby the signal from the switches 24 and 24a when both oleo legs becomeyfully extended, subject to the time delay T from the start of phase Bas in the case of the FIGURE 2 embodiment. The output of the computer 88is fed to a director index pitch servo 90 and the output of the latterunit is used to actuate the director indicator index 91, shown in FIGURE4. The signal from the gyroscope 16 is changed from the rate to the rate+I rate form Iby a control'signal from the pitch demand computer 88applied to the gyroscope over a linemarked control line.y f

The pitch attitude signal 0 is applied to a pitch attitude display servounit 92 which in turn feeds the indicator instrument of FIGURE 4.v v

The signal (ip-tpD) k5 from the limiter 81 is applied l to a roll demandcomputer 93, which also receives Vthe bank angle signal qb from the unitA61 and the constant signal from the units 83 and 96 and feeds adirector index-roll servo 94, the output of which is also applied to thein-' dicator instrument of FIGURE 4.

9 e 10 The bank angle signal is also applied to a bank pitch rate.During this phase the aircraft continues to attitude display servo 95the `output of which is also accelerate and if AHR has been correctlychosen the airapplied to the indicator instrument of FIGURE 4. v craftwill leave the runway with an optimal nose-up The sequence and method ofoperation of the-system pitch rate of about 1 to 11/2/second, which isslightly are described below, and in order to promote a ready less thanthat required in the lirst part of the pull-up. understanding of thedescription, the switching modes 5 Phase C.-As described in relation toFIGURE 2, the

of the pitch demand computer and the form of output aircraft incidenceis now further increased to produce signal to the indicator instrumentfor the different phases upward acceleration. are given in tabular form,as follows: The phase is initiated by full extension of both port TABLEOutput of pitch demand Phase of tal-:eci Conditions for phase engagementPitch gyro Inode computer None engaged Power on Rate -kqq (when gyrorotor has run l1 A (pre-rotation) Power on, director on, V Va .dokvf(V-Vn)-kqq B (rotation) Power on, director on,V Viz Rate -l-f rate K6(A9R-fqdt)-k. q C (pull-up) and D (initial climb) Power on, director on,both main U/C oleos Rate kv (V VT0C) k3-kqq fully extended, and at leasttime T after phase B engaged.

Power OIL-When the power is switched on the pitch and starboard mainoleo legs closing microswitches, demand computer passes the gyro ratesignal to the initiation being however prohibited until at least time Tdirector and the index immediately centres under. the after the start ofPhase B; T will be 0f order 2 seconds. action of the electrical springon the gyro girnbal/rotor The pitch demand computer switches the pitchgyroscope system. When the rotor has run up, the index will back fromthe rate -l-f rate to the rate mode and passes move about the zerodemand marker with aircraft pitch the signal [kv(V'-VToC)] k3 through tothe indicator. rate, the upward displacement being -kqq. The pull-up isthus demanded by displacing the indi- Dzrector on.-Before or duringtaxiing to the take-oif 3o cator index upwards by an amount proportionalto starting point the director system is switched on, as in [kv(V-VToC)]k3-kqq. The initial pitch rate demand the case of the embodiment ofFIGURE 2, whereupon is normally ka/kq and is limited to some value oforder the pitch demand computer passes the kvf(V-VR) signal 11/2 /secondwhich, in a steady pull-up (incidence conthrough to lthe director,making the total signal to the stant) at a true airspeed of 120 knots,gives about 1/6 g. director now kvf(V-VR)-kqq, where the function Thelimiting value is determined by the constant k3. When f(V-VR) is equalto V-VR except for V VRW when the phase-advanced speed demand signal thevalue is limited to VRW- VR, the limit being imposed by the constant k4.The index is displaced downwardly kv(V-VToo) by an amount corrsepondingto the preset value of' VR VRW to the marker 53. 40 later comes off itslimit k3 the pilot is directed to take Phase A Tnis phase and the eventsand conditions off the pitch rate and this should be accomplished intime associated therewith, are identical with those described gfveme'flby the um@ Constant of Phase-advance 0f V'- in relation to theembodiment of FIGURE 2 and no Since this pitch rate of l1/2 /second isonly slightly greater` further descrinnon'is deemed necessary thedesigned-for value at lift-off the index should As in Director on thesignal to the director is mmluy move sllghtly .abOV Centre by an aIHOuDPFO- kvf(V VR)-kqq portional to leg-knap The index should be central, orPhase B.A rapid and chosen increase of incidence is nearly S0 at. theend of Phase B, and hence il Will be now rednned-` expected to make aninsignicant movement (probably As the airspeed reaches VR the change ofSign of the upwards) on engagement of Phase C. Thus, although the fg/'/R) Signal into the nnen demand computer Simn1 control laws 1n Phase Band Phase difer considerably taneously switches out this signal andswitches the gyrothis s hould not be appafent to llledallo, WllO WlllrOtate scope 16 from me rate to the rate H rate mode Sothat the lift-offand pull-up into the initial climb as a single` the index, now displacedfrom centre by` an amount maneuverproportional to K(A0R-fq)-kqq, flicksup to demand Phase D Alhoogn no SWlolllng aenon oeellrS lm he nose nnrotation of amount 0R at an initial rate computer after initiation ofphase C, this initial climb phase begins when the airspeed is withinsome small- MR margin (say 2 or 3 knots) of VTOC (by which time the lcLlundercarriage should be substantially retracted) and con- This pitchrate demand then falls oif and by the time the tlnues to the Start ofthe ell-foute Climb at llefllallsiV index is steady at centre with themean pitch rate zero li'500feetl1e1gl1 the aircraft ywill have rotatedup'stably through the angle 60 The norfnal obleotlve S .to hold theairspeed at the MR. The value of AHR is, as previously stated, chosenA"aloe vs WhlCh tyPlCally eXCeedS V2 by at leaS l0 kIlOfS, somewhatgreater than the value which will give lift-olf, and PoSSlbly 30 o1' 40knots 0r more but lf, abnormally,

with me object of obtaining 1ift off with a pitch rate. the value of his below a certain value k2 then the de-v slightly less than thatrequired in the early part of the mondeo Speed 1S reduced by an amountPropoltlonal by.

nnu nn with the proviso that the pnenrate at nfaoff the preset constant`k1 to this decit but with a lower must not be so great as to result inthe aircraft tail-skid limit of V2 When h'S the Pl'eset Constant Valuek2 In touching the runway f special circumstances, such as lack of poweror deficient The maximum director index displacement may bePerfolffnaneei the available p oWel SllOUld be Used for made toconesnond to about 15 of pitch angle demand climbing rather thanaccelerating. Although h may diifei with k,=3k, this maximumdisplacement wiiiaiso cory7o Somewhat from rete 0f @limb h it may besaidthatfor 1.esnond to 5 /Seoond nitoh rate demand Sotaking a nearly'constant climb rate the two quantities will diifei' typical A0B of 12,the index 91 will initially move' fourv little and that the value k2 isclosely related to the .climb fifths of the distance to the top of theYscale atr-the start rate aPPfoPflatefo'alfSPeed V2 ln theScheduled-takeof Phase B onto the marker 59 shown in FIGURE 4 and offnet climb path.

this may also vbe considered as a demand for a 4/second 75 The value ofthe takeoff climb speed defined in this The speeds V2 and V3 are in thenature of variable constants to the extent that they are generated asfunctions the of both aircraft weight and flap angle, the former being7Uset in by the pilot and the latter sensed automatically, as explainedpreviously. The ap angle is sometimes reduced quite soon after takeoff(but not below 400 feet under present civil aircraft regulations) andthe value of VTOC will then usually increase as the flaps areretracting, due

to the consequent increases in V2 and V3; these increases will, however,be small, of the order knots. The director index will move appropriatelyto demand VTOC at every instant, allowing also for the eifects of ilapangle change on V' and h' and of any trim change at constant stickposition and airspeed.

The system operation in takeoii` and initial climb with wings level hasbeen described without consideration of either thrust loss orapplication of noise abatement procedures. Nor has the way in which thesystem directs the pilot to carry out baulked landing procedure-beenconsidered. The operation of the system under these conditions will nowbe described. Lateral control of the aircraft will not be specicallydescribed since this is already well known.

EFFECT OF THRUST LOSS The system is designed to give optimal ornear-optimal short and long term performance in this important, thoughrare, eventuality.

Thrust loss during rotation or just after lift-ot will have no immediateeffect on the indicator though the pull-up will be of shorter durationbecause the term kv(V'-VTOC) will come oi its limit k3 sooner due t0 thereduction of V in spite of the possible reduction in Vroc- Later thrustloss may result in a more or less rapid downward movement of thedirector index and the answering pilots forward stick movement willcheck the reduction in airspeed which might otherwise occur.

The steady climb will be at a atter angle than normal and usually at alower airspeed, possibly as low as V2, but exceeding V2 if the climbrate is sufficiently high despite the subnormal thrust.

NOISE ABATEMENT Since noise abatement procedure must sometimes befollowed it is necessary to make provision for changing one or moreconstants of the system, for example k1 and/ or k2, as above described,via the pilot-*operated noiseabate switch 69 which will be operated insympathy with the thrust reduction. s

. BAULKED LANDING lf on 'the landing approach the estimated aircraftweight is set up in the director system, then with power on and directoron and undercarriage down the system will be operating in phase C(pull-up) since the airspeed will surely exceed the VR appropriate tothis weight. The

speed VTOC may increase from V2 towards V3 which may reducev or eveneliminate the speed drop otherwise necessary to get to VTOC. 5

EXTRA PITCH RATE DEMAND IN TURNS Neglecting the effect of climb angle,which is smaller for the type of aircraft likely to use the system, theextra pitch rate demand in radians/ second is g/ VT sin p tan e, whereVT is the true airspeed. This is shown in FIGURE 3 approximated by g/ V2sin qb tan qS Y where V2 is C.A.S. (corrected airspeed) or possiblyI.A.S.

(indicated airspeed).

- Due to this the pitch signal to the director, in phases C and D, i.e.,when airborne, is v v In the present invention, the leads 111, 112 and113 shown in FIG. 3, instead of being connected directly t0 the computer86 are connected as shown in FIG. 5, in order to prevent the demandedclimb speed from decreasing.

The lead 111, which carries signal V2, is connected as one input of asummation amplifier 114 whose other input is connected to the output ofa multiplication device 115 to the inputs of which are connected theleads 112 and 113, carrying respectively signal K1, and signal (h-K2).

The output from summation amplifier 114, which is representative of ademanded climb speed (V'Toc), is fed as one input to a further summationampliiier 116. The output from the summation amplifier 116 is fed via adiode D1 as the input to an integrator 117. The output from theintegrator 117 is fed back via an inverter amplifier 118 as a secondinput to the summation amplifier 116, the output from integrator 117also being utilised as the output of the circuit.

In operation the computed demand climb speed V'TOC is fed via sum-mationamplifier 116, diode D1 and integrator 117 to the output of thearrangement, the out-4 put of integrator 117 being inverted and fed backas the other input to summation amplifier 116. Hence so long as thedemanded climb speed signal V'TOC is increasing the output of thearrangement VTOC continues to increase. However when the demanded climbsignal VTOC decreases yfor any reason the fed back signal -VT0C becomesthe greater and the output from summation arnplier 116, being negative,is blocked by the diode D1. The output from integrator 117 maintains itssteady state value and hence the output signal Vfl-OC does not decrease.Hence it will be seen that the demanded climb speed VTOC fed to thedirector can thus only increase and not decrease, so preventing a lowerspeed being demandedinthe event of an engine failure away from theground Y,when the aircraft is ying at a speed significantly above its V2speed. t Y The integrator 117 may be any common type of integrator usedin aircraft ilight control systems, e.g. either an electro mechanicalyintegrating, motor driving potentiometers or syncro pick-offs or a pureelectronic integrating device.

I claim: 1. A system for producing a demanded climb speed signal VTOCfor an aircraft during the. initial climb cycle y of its takeoff,comprising, in. combination,

means for producing a computed demand climb speed signalV1ToC=V2+K1(h-K2) where: w

.V2 is minimum climb out speed, l h israte of climb, and Y K1 and K2 areconstants, integrating means having an input and said demanded climbspeed signal VTOC as its output,

means for inverting the output of said integrating means,

summing means for combining the inverted signal VTOC and said signalVTOC as the input of said integrator means,

and means for -blocking the input to said integrator means when VTOCV1TOC so that the demanded climb speed signal VTOC may not decreaseduring the initial climb cycle.

2. In the system as dened in clam 1, wherein the last mentioned meanscomprises a one-way device connecting the combined signal to saidintegrating means.

3. In the system deiined in claim 2 wherein said one- Way device is adiode.

1 4 References Cited UNITED STATES PATENTS 3,295,369 1/1967 Priestley73-178 3,241,362 3/1966 Scott 73-178 FOREIGN PATENTS 965,308 7/1964Great Britain.

LOUIS R. PRINCE, Primary Examiner DONALD O. WOODIEL, Assistant Examiner

